The present invention relates to space satellite systems architecture, and more particularly, to a multi-mission adaptable modular spacecraft bus having improved structural, thermal, and accessibility characteristics and reduced part count.
In the past, almost every new space program would develop a spacecraft bus design optimized for a specific space mission. Previous common bus designs failed to allow for ease in tailoring the bus to accommodate mission differences. This design approach has lead to a proliferation of mission specific bus designs.
For cost and performance effectiveness in today's competitive environment, satellite buses must be (1) scalable for various missions, (2) adaptable to fit a variety of launch vehicles, and (3) provide economies of scale in development and production cycles through the use of commercial subsystems and components and manufacturing and testing processes.
Current state of the art spacecraft fabrication and assembly techniques use a high degree of composites. Composites offer many advantages over the metallic materials commonly used in the primary supporting structure of conventional spacecraft designs. A key benefit of composites is that they can provide significant weight reductions in the final product. Composites are more easily fashioned into structural pieces of complex geometry (e.g. rounded surfaces. irregular profiles, etc.), and thus extensive use of composites can help to reduce part count as well as reduce the number of mechanical fasteners that are required to secure the parts or structural pieces together.
In accordance with modular satellite assembly techniques, the major structural components are assembled from composite parts and bolted together, including a series of access panels. In order to reduce weight, composites are used for the spacecraft. Although such state of the art spacecraft designs have made great strides in reducing part count and weight, the use of composites in these spacecraft designs still relies heavily on geometry and joint methods from traditional metal fabrication and assembly techniques. Areas of improvement include using the composites in a way to maximize the benefits composites offer over conventional fabrication materials in order to further reduce total part count.
Studies have shown that the number of parts is directly proportional to assembly cost increases. For each part, there must be a designer, a checker, a planner and planning paper, an expeditor or subcontract manager and more paper, etc. In addition, more parts in an assembly complicates the assembly process, which results in more assembly tooling and assembly time. Reducing part count is critical to reducing assembly costs.
On the other hand, if part reduction goes to the extreme and results in only a few parts that are extremely complex, the end result can be high part scrap and rework rate, and long fabrication times, thus negating the cost saving of minimizing part count. One of the design goals of the present invention is to strike a balance between minimum part count and part complexity.
It is well understood that the efficiency of composites is decreased as the number of fasteners and discontinuous joints are increased in the completed structure. To obtain the maximum benefits of composites, spacecraft designers must rethink the way they fabricate and assemble the major structural components of spacecraft. A continuous primary structure having a minimum number of bolt together fasteners would be a more efficient use of composites. Of course, a continuous composite containment structure for satellite primary structure fabrication is of little benefit if it does not permit good access to the spacecraft interior for installation and testing of the spacecraft subsystems prior to launch.
An example of a typical prior art modular spacecraft using a high degree of composites is designated generally by reference numeral 10 in FIG. 1. This configuration, which includes a payload module 14 and core module 16 with integral propulsion subsystem 18, is specifically designed as a geosynchronous communications satellite and is optimized for high altitude staring payloads. The typical prior art modular spacecraft 10 is characterized by a relatively high part count, low stiffness, and relatively low thermal and dimensional stability and therefore is not suitable for low altitude missions, such as remote sensing missions, that use agile pointing payloads.
Inadequate or limited access to the spacecraft interior is also a problem associated with most state of the art modular spacecraft. A case in point is the procedure that is required for installing the integral propulsion subsystem 18 in the typical prior art modular spacecraft 10. In view of the limited access to the spacecraft interior, and further in view of the presence of other pre-installed subsystems, the propulsion subsystem 18 must be installed as a number of subassemblies, each of size small enough to fit within the access panel openings. Once the propulsion system subassemblies are inside the access panels, they must be maneuvered around the other pre-installed subsystems into their assigned locations. The various subassemblies of the propulsion subsystem are then welded together in place. Field welding of this nature is both costly and time intensive since it must be done in a clean room environment and it further requires use of special portable welding apparatus so as not to compromise the other subsystems of the spacecraft.
Much greater manufacturing and assembling efficiencies could be realized if the propulsion subsystem could be installed within the spacecraft as a fully assembled plug and play unit. Accordingly, a modular spacecraft design having reduced part count and number of fasteners and a structural geometry that enables full and unimpeded access to the various subsystems of the spacecraft during assembly, installation and testing would constitute a significant advancement.
The typical prior art spacecraft 10 also does not package well in that it does not allow for growth by adding additional subsystem components such as electronics boxes, reaction wheels, etc, without further compromising the structural stiffness and stability of the spacecraft. As noted above, a number of the spacecraft subsystems, such as the propulsion subsystem, must be installed as a number of subassemblies. This design does not permit easy and convenient subsystem removal for repair, replacement and/or upgrade.
Thus, it would be desirable to provide a modular spacecraft structure in which the various subsystems are segregated into separate modules to permit parallel production and testing. A further improvement in the manufacture of spacecraft would be the provision of a standard bus that can also be fabricated in advance and in an efficient time frame. Further still, there is a need for a modular spacecraft bus that is easily produced and is easily adaptable and scalable to a wide range of satellite missions.